In a typical gas turbine engine, energy is exchanged between an axial flow of working fluid and a rotating shaft by conducting the fluid through a series of turbine stages. Each stage consists of a rotor comprised of a plurality of blades disposed on the shaft and, just upstream of each rotor, a stator comprised of a plurality of stationary blades or vanes. The purpose of the stator is to optimize the velocity profile for the working fluid entering the rotor stage in order to improve energy transfer efficiency at the rotor and thus improve overall engine efficiency.
The primary source of energy is typically a combustion process which raises the temperature of the working fluid in direct proportion to the energy added by the combustion process. Even though turbine inlet temperatures of 3600.degree. F. (2000.degree. C.) are possible, the temperature of the working fluid is limited by the ability to maintain the vane metal temperatures at acceptable levels, especially in the first stages.
One solution is to install a cooling system in the turbine vanes. The typical, prior art cooling system routes a portion of the relatively cool compressor bleed air, which otherwise would be used in the combustion process, into cavities in the vanes. This cooling air then passes through an array of cooling holes in the surface of the vane and into the flow of the working fluid. The density of cooling holes increases towards the leading edge of the vane where the external heat transfer rate is highest for each vane. The cooling air transfers heat from the vane during the passage through the cavity and cooling holes and creates a film of cooling air which flows over the external surface of the vane. Ideally this film of cooling air mixes gradually with the working fluid as it flows down the surface of the vane and provides a buffer layer between the hot working fluid and the vane surface. In practice, however, the momentum of the cooling air passing through the holes causes mixing of the cooling air and the working fluid at the vane surface reducing the effectiveness and increasing the number of cooling holes required.
Unfortunately, there are significant drawbacks to this type of cooling system. First, the injection of cooling air through the holes increases the thickness of the boundary layer on the vane surface and increases the aerodynamic loss. Second, the overall efficiency of the engine suffers due to having to divert a portion of compressed air away from the combustion process. Third, the creation of these arrays of holes in the vanes adds to the fabrication costs of each vane and to the total fabrication costs of a turbine. Fourth, the array of holes produces local temperature gradients across the surface of the vane and induces thermal stresses in the vane which may eventually degrade the vane surface. Finally, the performance of the cooling system degrades with time if it is utilized in a particle contaminated environment, such as is the case for aircraft applications, due to direct impingement of foreign particles on the cooling holes.
Another possible solution is to fabricate the vane entirely from a material capable of withstanding high temperatures, such as a high strength ceramic as suggested in U.S. Pat. No. 4,768,924. A properly designed ceramic vane would be able to withstand the high temperature without requiring a cooling system. Such vanes fabricated completely from ceramics are not currently practical. The primary limiting factors are the catastrophic failure mode of the ceramic material, the tendency of the failure to propagate rapidly throughout the ceramic components of the turbine, and the differing rates of thermal expansion between the ceramic components and the non-ceramic supporting components.